CHAPTER 3
Drag Force and Drag Coefficient
From: Sadraey M., Aircraft Performance Analysis, VDM Verlag Dr. Müller, 2009
3.1. Introduction
Drag is the enemy of flight and its cost. In chapter 2, major forces that are influencing
aircraft motions were briefly introduced. One group of those forces is aerodynamic forces
that split into two forces: Lift force or lift, and Drag force or drag. A pre-requisite to
aircraft performance analysis is the ability to calculate the aircraft drag at various flight
conditions. One of the jobs of a performance engineer is to determine drag force
produced by an aircraft at different altitudes, speeds and configurations. This is not an
easy task, since; this force is a function of several parameters including aircraft
configuration and components. As it was discussed in chapter 2, the drag is a function of
aircraft speed, wing area, air density, and its configuration. Each aircraft is designed with
a unique configuration, thus, aircraft performance analysis must take into account this
configuration. The configuration effect of aircraft drag is represented through the drag
coefficient (CD), plus a reference area that relates to the aircraft.
An aircraft is a complicated three-dimensional vehicle, but for simplicity in
calculation, we assume that the drag is a function a two-dimensional area and we call it
the reference area. This area could be any area including tail area, wing area and fuselage
cross sectional area (i.e., fuselage cross section), fuselage surface area, and even aircraft
top-view area. No matter what area is selected, the drag force must be the same. This
unique drag comes from the fact that the drag coefficient is a function of the reference
area. Therefore, if we select a small reference area, the drag coefficient shall be large, but
if we choose a large reference area, the drag coefficient shall be small. In an air vehicle
with a small wing area (e.g., high-speed missile), the fuselage cross-sectional area
(normal to the flow) is often considered as the reference area. However, in an aircraft
with a large wing, the top-view of wing; planform area (in fact gross wing area) is often
assumed to be the reference area.
The measurement of this area is easy; and it usually includes the most important
aerodynamic part of the aircraft. This simplified reference area is compensated with the
complicated drag coefficient, as we discussed in chapter 2.
Aircraft Performance Analysis
1
D
1
V 2 SC D
2
(3.1)
The drag coefficient (CD) is a non-dimensional parameter, but it takes into
account every aerodynamic configuration aspect of the aircraft including large
components as wing, tail, fuselage engine, and landing gear; and small elements such as
rivets and antenna. This coefficient has two main parts (as will be explained in the next
section). The first part is referred to as lift-related drag coefficient or induced drag
coefficient (CDi) and the second part is called zero-lift drag coefficient (CDo). The
calculation of the first one is not very hard, but it takes a long time and energy to
determine the second part. In large transport aircraft, this task is done by a group of
engineers up to twenty engineers for a time period of up to six months. For this reason, a
large portion of this chapter is devoted to the calculation of CDo. This calculation is not
only time consuming, but also is very sensitive, since it influences every aspect of aircraft
performance.
One of the occasions in which the drag is considered a beneficial factor and is
effectively used is in parachute. A parachute is a device employed to considerably slow
the motion of an object/vehicle through an atmosphere (e.g., Earth or Mars) by increasing
drag. Parachutes are used with a variety of loads, including people, food, equipment, and
space capsules. Drogue chutes are used to sometimes provide horizontal deceleration of a
vehicle (e.g., space shuttle after a touchdown). The parachute is utilized by paratroopers
to extremely reduce the terminal speed for a safe landing.
One of the primary functions of aerodynamicists and aircraft designers is to
reduce this coefficient. Aircraft designers are very sensitive about this coefficient,
because any change in the external configuration of aircraft will change this coefficient
and finally aircraft direct operating cost. As a performance engineer, you must be able to
estimate the CDo of any aircraft just by looking at its three-view with an accuracy of about
30%. As you spend more time for calculation, this estimation will be more accurate, but
will never be exact, unless you use an aircraft model in a wind tunnel or flight test
measurements with real aircraft model. The method presented in this chapter is about
90% accurate for subsonic aircraft and 85% for supersonic aircraft.
3.2. Drag Classification
Drag force is the summation of all forces that resist against aircraft motion. The
calculation of the drag of a complete aircraft is a difficult and challenging task, even for
the simplest configurations. We will consider the separate sources of drag that contribute
to the total drag of an aircraft. The variation of drag force as a function of airspeed looks
like a graph of parabola. This indicates that the drag initially reduces with airspeed, and
then increases as the airspeed increases. It demonstrates that there are some parameters
that will decrease drag as the velocity increases; and there are some other parameters that
will increase drag as the velocity increases. This observation shows us a nice direction for
drag classification. Although the drag and the drag coefficient can be expressed in a
number of ways, for reasons of simplicity and clarity, the parabolic drag polar will be
Aircraft Performance Analysis
2
used in all main analyses. Different references and textbooks use different terminology,
so it may confuse students and engineers. In this section, a list of definitions of various
types of drag is presented, and then a classification of all of these drag forces is
described.
Induced Drag: The drag that results from the generation of a trailing vortex system
downstream of a lifting surface with a finite aspect ratio. In another word, this type of
drag is induced by the lift force.
Parasite Drag: The total drag of an airplane minus the induced drag. Thus, it is the drag
not directly associated with the production of lift. The parasite drag is composed of drag
of various aerodynamic components; the definitions of which follow.
Skin Friction Drag: The drag on a body resulting from viscous shearing stresses (i.e.,
friction) over its contact surface (i.e., skin). The drag of a very streamlined shape such as
a thin, flat plate is frequently expressed in terms of a skin friction drag. This drag is a
function of Reynolds number. There are mainly two cases where the flow in the boundary
layer is entirely laminar or entirely turbulent over the plate. The Reynolds number is
based on the total length of the object in the direction of the velocity. In a usual
application, the boundary layer is normally laminar near the leading edge of the object
undergoing transition to a turbulent layer at some distance back along the surface.
A laminar boundary layer begins to develop at the leading edge and its thickness
grows in downstream. At some distance from the leading edge the laminar boundary
becomes unstable and is unable to suppress disturbances imposed on it by surface
roughness or fluctuations in the free stream. In a distance the boundary layer usually
undergoes a transition to a turbulent boundary layer. The layer suddenly increases in
thickness and is characterized by a mean velocity profile on which a random fluctuating
velocity component is superimposed. The distance, from the leading edge of the object to
the transition point can be calculated from the transition Reynolds number. Skin friction
factor is independent of surface roughness in laminar flow, but is a strong function of
surface roughness in turbulent flow due to boundary layer.
Form Drag (sometimes called Pressure Drag): The drag on a body resulting from the
integrated effect of the static pressure acting normal to its surface resolved in the drag
direction. Unlike the skin friction drag that results from viscous shearing forces tangential
to a body’s surface, form drag results from the distribution of pressure normal to the
body’s surface. In an extreme case of a flat plate normal to the flow, the drag is totally the
result of an imbalance in the pressure distribution. As with skin friction drag, form drag is
generally dependent on Reynolds number. Form drag is based on the projected frontal
area. As a body begins to move through the air, the vorticity in the boundary layer is shed
from the upper and lower surfaces to form two vortices of opposite rotation. A number of
symmetrical shapes having drag values [13] at low speed are illustrated in Table 3.1. The
drag coefficient values in this table are based on the frontal area. In this table, the flow is
coming from left to the right.
Aircraft Performance Analysis
3
No
1
Body
Square rod
2
Circular rod
3
Equilateral
triangular rod
4
Rectangular rod
Status
Sharp corner
Elliptical rod
1.2
Laminar flow
Turbulent flow
Sharp edge face
1.2
0.3
1.5
Flat face
2
Sharp corner
Laminar flow
Turbulent flow
6
7
Symmetrical shell
Semicircular rod
CD
2.2
Round corner
Round front
edge
5
Shape
L/D = 0.1
L/D = 0.5
L/D = 3
L/D = 0.5
L/D = 1
L/D = 4
L/D = 2
L/D = 8
L/D = 2
L/D = 8
1.9
2.5
1.3
1.2
0.9
0.7
0.6
0.25
0.2
0.1
Concave face
2.3
Convex face
1.2
Concave face
1.2
Flat face
1.7
a. Two dimensional bodies (L: length along flow, D; length perpendicular to the flow)
Body
Laminar/turbulent
Status
CD
Cube
Re > 10,000
1.05
Thin circular disk
Re > 10,000
1.1
Re > 10,000
0.5
Cone ( = 30o)
5
Sphere
Laminar Re < 2×10
0.5
Turbulent Re > 2×106
0.2
5
Ellipsoid
Laminar Re < 2×105
0.3-0.5
Turbulent Re > 2×106
0.1-0.2
6
Hemisphere
Re > 10,000
Concave face
0.4
Re > 10,000
Flat face
1.2
7
Rectangular plate
Re > 10,000
Normal to the flow
1.1 - 1.3
5
8
Vertical cylinder
Re < 2×10
L/D = 1
0.6
L/D = ∞
1.2
9
Horizontal cylinder
Re > 10,000
L/D = 0.5
1.1
L/D = 8
1
10 Parachute
Laminar flow
1.3
No
1
2
3
4
b. Three dimensional bodies (L: length, D; diameter)
Aircraft Performance Analysis
4
Table 3.1. Drag coefficient values for various geometries and shapes
Interference Drag: The increment in drag resulting from bringing two bodies in
proximity to each other. For example, the total drag of a wing-fuselage combination will
usually be greater than the sum of the wing drag and fuselage drag independent of each
other.
Trim Drag: The increment in drag resulting from the (tail) aerodynamic forces required
to trim the aircraft about its center of gravity. Trim drag usually is a form of induced and
form drag on the horizontal tail.
Profile Drag: Usually taken to mean the total of the skin friction drag and form drag for
a two-dimensional airfoil section.
Cooling Drag: The drag resulting from the momentum lost by the air that passes through
the power plant installation for the purpose of cooling the engine.
Wave Drag: This drag; limited to supersonic flow; is a form of induced drag resulting
from non-canceling static pressure components to either side of a shock wave acting on
the surface of the body from which the wave is emanating.
Drag
Zero-lift Drag
Induced Drag
Miscellaneous
Drag
Wave
Drag
Skin friction
Drag
Nacelle
Wing
Fuselage
Interference
Drag
Trim
Drag
Cooling
Drag
Form
Drag
Tail
Landing
gear
Compressibility
Drag
Vortex
Drag
Wave Drag
Strut
CL
dependent
Volume
dependent
Figure 3.1. Drag classification
The material to follow will consider these various types of drag in detail and will present
methods of reasonably estimating their magnitudes. Figure 3.1 illustrates the drag
classification into two major groups.
For a conventional aircraft, the drag is divided into two main parts; lift related
drag, and non-lift related drag. The first part is called induced drag (Di), because this drag
is induced by lift (pressure). The second part is referred to as zero-lift drag (Do), since it
does not have any influence from lift, and is mainly originates from shear stress.
Aircraft Performance Analysis
5
D  Do  Di
(3.2)
a. Induced drag: The induced drag is the drag directly associated with the
production of lift. This results from the dependency of the induced drag on the
angle of attack. As the angle of attack of the aircraft (i.e., lift coefficient) varies,
this type of drag is changed. The induced drag in itself consists of two parts. The
first part originates from vortices around wing, tail, fuselage, and other
components. The second part is because of air compressibility effect. In low
subsonic flight, it is negligible, but is high subsonic and transonic flight, must be
taken into account. In supersonic flight, wave drag (Dw) is added to the induced
drag. The reason is to account for the contribution of shock wave. The wing is the
major aircraft component contributor for the lift production. Thus, about 80% of
the induced drag comes from wing; about 10% comes from tail; and the rest
originate from other components. The induced drag is a function of airspeed, air
density, reference area, and the lift coefficient:
1
(3.3)
Di  V 2 SC D i
2
In this equation, the coefficient CDi is called induced drag coefficient. The method
to calculate this coefficient will be introduced in the next section. Figure 3.2
shows the behavior of induced drag as a function of airspeed. As the airspeed
increases, the induced drag decreases; so the induced drag is inversely a function
of airspeed.
Figure 3.2. Variations of Do and Di versus velocity
b. Zero-lift drag: The zero-lift drag includes all types of drag that do not depend on
production of the lift. Every aerodynamic component of aircraft (i.e. the
Aircraft Performance Analysis
6
components that are in direct contact with flow) generates zero-lift drag. Typical
components are wing, horizontal tail, vertical tail, fuselage, landing gear, antenna,
engine nacelle, and strut. The zero-lift drag is a function of airspeed, air density,
reference area, and the external shape of the components:
Do 
1
V 2 SC D o
2
(3.4)
In this equation, the coefficient CDo is called zero-lift drag coefficient. The
method to calculate this coefficient will be introduced in Section 3.4. Figure 3.2
shows the variation of zero-lift drag as a function of airspeed. As the airspeed
increases, the induced increases too; so the zero-lift drag is directly a function of
airspeed.
From the equations 3.1, 3.2, 3.3 and 3.4; one can conclude that drag coefficient has two
components:
CD  CD o  CD i
(3.5)
The calculation of CDi is not a big deal and will be explained in the next section;
but the calculation of CDo is very challenging, tedious, and difficult. Major portion of this
chapter is devoted to calculation of CDo. In fact, the main idea behind this chapter is about
calculation of CDo.
3.3. Drag Polar
The aircraft drag may be mathematically modeled by variety of methods. It seems natural
to seek the similarity of variation of drag due to a flight parameter to a standard figure or
geometry. We are looking for an accurate, but simple mathematical model, and a math
expression for such curves as in figure 3.3.
As figure 3.3 and equation 3.5 show, the drag is composed of two terms, one
proportional to the square of airspeed (i.e., V2) and the other one inversely proportional to
V2. The first term, called zero-lift drag represents the aerodynamic cleanness with respect
to frictional characteristics, and shape and protuberances such as cockpit, antennae, or
external fuel tanks. It increases with the aircraft velocity and is the main factor in
determining the aircraft maximum speed. The second term represents induced drag (drag
due to the lift). Its contribution is highest at low velocities, and it decreases with
increasing flight velocities. If we combine (indeed add) these two curves (Di and Do) in
Figure 3.2, we will have a parabolic curve such as what shown in figure 3.3. The
parabolic drag model is not exact; but accurate enough for the purpose of performance
calculation. A similar behavior is observed for the variation of drag coefficient versus lift
coefficient. Drag polar is a math model for the variation of drag coefficient as a function
of lift coefficient.
Aircraft Performance Analysis
7
Figure 3.3. Variations of drag versus airspeed
Although the drag and the drag coefficient can be expressed in a number of ways, for
reasons of simplicity and clarity, the parabolic drag polar will has been selected in the
analysis. This is true only for subsonic flight. For the existing supersonic aircraft, the
drag cannot be adequately described by such a simplified expression. Exact calculations
must be carried out using extended equations or tabular data. However, the inclusion of
more precise expressions for drag at this stage will not greatly enhance basic
understanding of performance, and thus, will be included only in some calculated
examples and exercises. Note that the curve begins from stall speed, since an aircraft is
not able to maintain a sustained level flight at any speed below the stall speed. The same
conclusion is true for the variation of drag coefficient (CD) versus lift coefficient (CL) as
shown in figure 3.4.
A non-dimensionalized form of figure 3.3; the variation of drag coefficient versus
lift coefficient; is demonstrated in figure 3.4. It can be proved that a second order
parabolic curve can mathematically describe such a curve with an acceptable accuracy:
y  a  bx 2
(3.6)
where y may be replaced with CD and x may be replaced with CL. Therefore, drag
coefficient versus lift coefficient is modeled with the following parabolic model:
C D  a  bC L
2
(3.7)
Now, we need to determine the values or expressions for “a” and “b” in this
equation. In a symmetrical parabolic curve, the parameter “a” is the minimum value for
parameter “y”. hence, in a parabolic curve of CD versus CL, the parameter “a” must be the
minimum amount of drag coefficient (CDmin). We refer this minimum value of drag
Aircraft Performance Analysis
8
coefficient as CDo as it means the value of CD when the lift is zero. The corresponding
value for “b” in equation 3.7 must be found through experiment. Aerodynamicists have
represented this parameter with the symbol of “K”, and refer to it as induced drag
correction factor. The induced drag correction factor is inversely proportional to the wing
aspect ratio (AR) and wing Oswald efficiency factor (e). The mathematical relationship is
as follows:
CD
CDo
CL
Figure 3.4. A typical variations of CD versus CL
K
1
  e  AR
(3.8)
The wing aspect ratio is the ratio between the wing span (b) and the mean aerodynamic
chord (MAC or C). The ratio can be reformatted to be a function of the wing area (S) and
wing span as follows:
b bb b 2
AR  

C Cb
S
(3.9)
The wing Oswald efficiency factor represents the efficiency of a wing in
producing lift, and is a function of wing aspect ratio and the leading edge sweep angle,
LE. If the lift distribution is parabolic, the Oswald efficiency factor is assumed to be the
highest (i.e., 100% or 1). The Oswald efficiency factor is usually between 0.7 and 0.9.
Ref. 9 introduces the following two expressions for estimation of Oswald efficiency
factor:


e  4.61 1  0.045 AR 0.68 cos LE 

e  1.78 1  0.045 AR
0.68
Aircraft Performance Analysis
 0.64
0.15
 3.1
(3.10a)
(3.10b)
9
Equation 3.10a is for swept wings with leading edge sweep angles of more than
30 degrees and equation 3.10b is for rectangular wings (without sweep). These two
formulas are only valid for wing with high aspect ratio (e.g., AR more than 6).
x
y
Top view
Figure 3.5. Wing leading edge sweep angle
The wing leading edge sweep angle (see figure 3.5) is the angle between wing
leading edge and the aircraft y-axis. Table 3.2 shows wing Oswald efficiency factor for
several aircraft. The value of "e" is decreased at high angle of attacks (i.e., low speed) up
to about 30%.
Figure 3.6. The glider Schleicher ASW22 with wing span of 25 meter and CDo of 0.016
Aircraft Performance Analysis
10
Employing the induced drag correction factor (K), we have a mathematical expression for
the variation of drag coefficient versus lift coefficient.
C D  C Do  KC L
2
(3.11)
This equation is sometimes referred to as aircraft “drag polar”. The main
challenge in this equation is the calculation of zero-lift drag coefficient. Table 3.2 shows
typical values of CDo for several aircraft. The values in this table are the lowest values;
which means at the lowest airspeed (usually low subsonic speeds). Gliders or sailplanes
are aerodynamically the most efficient aircraft (with CDo as low as 0.01) and agricultural
aircraft are aerodynamically the least efficient aircraft (with CDo as high as 0.08) .The lift
coefficient is readily found from equation 2.3. Compare the glider Schleicher ASW22
that has a CDo of 0.016 (Figure 3.6) with the agricultural aircraft Dromader, PZL M-18
that has a CDo of 0.058 (figure 3.7).
Figure 3.7. Agricultural aircraft Dromader, PZL M-18 with CDo of 0.058
Comparison between equations 3.5 and 3.11 yields the following relationship.
C Di  KC L
2
(3.12)
So, induced drag is proportional to the square of lift coefficient. Figure 3.4 shows the
effect of lift coefficient (induced drag) on drag coefficient.
3.4. Calculation of CDo
The equation 3.11 implies that the calculation of aerodynamic force of drag is dependent
on zero-lift drag coefficient (CDo). Since the performance analysis is based on aircraft
drag, the accuracy of aircraft performance analysis relies heavily on the calculation
accuracy of CDo. This section is devoted to the calculation of zero-lift drag coefficient and
is the most important section of this chapter. The method by which the zero-lift drag
coefficient is determined is called “build-up” technique.
Aircraft Performance Analysis
11
No
1
2
3
4
5
6
7
8
9
10
Aircraft type
Twin-engine piton prop
Large turboprop
Small GA with retractable landing gear
Small GA with fixed landing gear1
Agricultural aircraft with crop duster
Agricultural aircraft without crop duster
Subsonic jet
Supersonic jet
Glider
Remote controlled model aircraft
CDo
0.022-0.028
0.018-0.024
0.02-0.03
0.025-0.04
0.07-0.08
0.06-0.065
0.014-0.02
0.02-0.04
0.012-0.015
0.025-0.045
e
0.75-0.8
0.8-0.85
0.75-0.8
0.65-0.8
0.65-0.7
0.65-0.75
0.75-0.85
0.6-0.8
0.8-0.9
0.75-0.85
Table 3.2. Typical values of "CDo" and "e" for several aircraft
As figure 3.8 illustrates, external aerodynamic components of an aircraft are all
contributing to aircraft drag. Although only wing; and to some extents, tail; have
aerodynamic function (i.e., to produce lift), but every component (either large size such
as the wing or small size such as a rivet) that has direct contact with air flow, is doing
some types of aerodynamic functions (i.e., producing drag). Thus, in order to calculate
zero-lift drag coefficient of an aircraft, we must include every contributing item. The C Do
of an aircraft is simply the summation of CDo of all contributing components.
Horizontal
tail
Vertical tail
Wing
Vertical tail
Landing Gear
Nacelle
Fuselage
Wing
Horizontal
tail
Figure 3.8. Major components of Boeing 737 contributing to CDo
1
This also refers to a small GA with retractable landing gear during take-off
Aircraft Performance Analysis
12
CDo  CDo f  CDow  CDoht  CDovt  CDoLG  CDoN  CDoS  CDoHLD  ...
(3.13)
where CDof, CDow, CDoht, CDovt, CDoLG, CDoN, CDoS, CDoHLD, are respectively representing
fuselage, wing, horizontal tail, vertical tail, landing gear, nacelle, strut, high lift device
(such as flap) contributions in aircraft CDo. The three dots at the end of equation 3.13
illustrates that there are other components that are not shown here. They include nonsignificant components such as antenna, pitot tube, stall horn, wires, interference, and
wiper.
Every component has a positive contribution, and no component has negative
component. In majority of conventional aircraft, wing and fuselage are each contributing
about 30%-40% (totally 60%-80%) to aircraft CDo. All other components are contributing
about 20%-40% to CDo of an aircraft. In some aircraft (e.g., hang gliders), there is no
fuselage, so it does not have any contribution in CDo; instead the human pilot plays a
similar role to fuselage.
In each subsection of this section, a technique is introduced to calculate the
contribution of each component to CDo of an aircraft. The primary reference for these
techniques and equations is Reference 1. Majority of the equations are based on flight test
data and wind tunnel test experiments, so the build-up technique is relying mainly on
empirical formulas.
3.4.1. Fuselage
The zero-lift drag coefficient of a fuselage is given by the following equation:
C Do f  C f f LD f M
S wet f
S
(3.14)
where, Cf is skin friction coefficient, and is a non-dimensional number. It is determined
based on the Prandtl relationship as follows:
Cf 
Cf 
0.455
log 10 Re2.58
1.327
Re
(Turbulent flow) (3.15a)
(Laminar flow) (3.15b)
The parameter Re is called Reynolds number and has a non-dimensional value. It is
defined as:
Re 
VL

Aircraft Performance Analysis
(3.16)
13
where is the air density, V is aircraft true airspeed,  is air viscosity, and L is the length
of the component in the direction of flight. For a fuselage, L is the fuselage length. For
lifting surfaces such as wing and tail, L is the mean aerodynamic chord.
Equation 3.15a is for a purely turbulent flow and equation 3.15b is for a purely
laminar flow. Most aircraft are frequently experiencing a combination of laminar and
turbulent flow over fuselage and other component. There are aerodynamic references
(e.g., [8] and [10]) that recommend technique to evaluate the ratio between laminar and
turbulent flow over any aerodynamic component. The transition point from laminar to
turbulent flow may be evaluated by these references. For simplicity, they are not
reproduced here. Instead, you are recommended to assume that the flow is either
completely laminar or completely turbulent. The assumption of complete turbulent flow
provides a better result; since over-estimation of drag is much better than its underestimation.
In theory, the flow is laminar when Reynolds number is below 4,000. However, in
practice, turbulence is not effective when Reynolds number is below 200,000; so when
the Reynolds number is less than 200,000, you may assume laminar flow. In addition,
when the Reynolds number is higher than 2,000,000, you may assume turbulent flow. As
a rule of thumb, in low subsonic flight, the flow is mostly laminar, but in high subsonic
and transonic speed, it becomes mostly turbulent. Supersonic and hypersonic flight
experiences a complete turbulent flow over every component of aircraft. A typical current
aircraft may have laminar flow over 10%-20% of the wing, fuselage and tail. A modern
aircraft such as the Piaggio 180 can have laminar flow over as much as 50% of the wing
and tails, and about 40%-50% of the fuselage.
The second parameter in equation 3.14 (fLD) is a function of fuselage length-todiameter ratio. It is defined as:
f LD  1 
60
L D 
3
L
 0.0025 
D
(3.17)
where L is the fuselage length and D is its maximum diameter. If the cross section of the
fuselage is not a circle, you need to find its equivalent diameter. The third parameter in
equation 3.14 (fM) is a function of Mach number (M). It is defined as
f M  1  0.08M 1.45
(3.18)
The last two parameters in equation 3.14 are Swetf and S, they are the wetted area of the
fuselage and the wing reference area respectively.
Wetted area is the actual surface area of the material making up the skin of the
airplane - it is the total surface area that is in actual contact with, i.e., wetted by, the air in
which the body is immersed. Indeed, the wetted surface area is the surface on which the
pressure and shear stress distributions are acting; hence it is a meaningful geometric
Aircraft Performance Analysis
14
quantity when one is discussing aerodynamic force. However, the wetted surface area is
not easily calculated, especially for complex body shapes.
Wing net area
Wing gross area
Figure 3.9. Wing gross area and wing net (exposed) area
A comment seems necessary regarding the reference area, S in equation 3.14. The
parameter S is nothing other than just a reference area, suitably chosen for the definition
of the force and moment coefficients.
The reference area (S) is simply an area as a basis or reference that can be
arbitrarily specified. This selection is primarily done for the convenience. The reference
area (S) for a conventional aircraft is the projected area that we see when we look down
on the wing from top view, including the fuselage section between two parts of the wing.
For this reason, for wings as well as entire airplanes, the wing planform area is usually
used as S in the definitions of CL, CD, and Cm. However, if we are considering the lift and
drag of a cone, or some other slender, missile like body, then the reference area, S is
frequently taken as the base area of the cone or fuselage. Figure 3.9 highlights the
difference between wing net area and wing gross area. Thus, when we say wing planform
area, we mean wing gross area. The reference area selection assumption will not incur
any error in the calculation of wing drag and aircraft drag. The reason is that the drag
coefficient will be automatically adjusted by this selection.
Whether we take the planform area, base area, or any other areas to a given body
shape for S, it is still a measure of the relative size of different bodies which are
geometrically similar. As long as you are consistent, you may take any significant area as
the reference area. What is important in the calculation of CL, CD, and Cm is to divide the
Aircraft Performance Analysis
15
aerodynamic force/moment to a noticeable area. It is urged to an aircraft performance
engineer that whenever you take a data for CL, CD, or Cm from the technical literature,
make certain that you know what geometric reference area was used for S in the
calculation. Then use that same area when making calculations involving lift, drag and
pitching moment. Otherwise, the results will involve significant inaccuracies. In contrast
with Swet, it is much easier to calculate the planform area of a wing.
3.4.2. Wing, Horizontal Tail, and Vertical Tail
Since wing, horizontal tail and vertical tail are three lifting surfaces, they are treated in a
similar manner. The zero-lift drag coefficient of the wing ( C Do ), horizontal tail ( C Do ),
w
ht
and vertical tail ( C Do ), are respectively given by the following equations:
vt
0.4
CDow
 S wetw
 C f w f tcw f M 
 S
 Cd min w 




 0.004 
CDoht
 S wetht
 C f ht f tcht f M 
 S
 Cd min ht 


 0.004 


 S wetvt
f M 
 S
 Cd min vt 


 0.004 


CDovt  C f vt f tcvt
(3.19)
0.4
(3.20)
0.4
(3.21)
In these equations, Cfw, Cfht, Cfvt are similar to what we defined for fuselage in formula
3.15. The only difference is that the equivalent value of L in Reynolds number (equation
3.22) for wing, horizontal tail, and vertical tail are their mean aerodynamic chords (MAC
or C ). In another word, the definition of Reynolds number for a lifting surface (e.g.,
wing) is:
Re 
V C

(3.22)
where the mean aerodynamic chord is calculated by
2 
 
(3.23)
C  C r 1   
3 
1   
where Cr denotes root chord (See figure 3.10), and  is the taper ratio; the ratio between
tip chord (Ct) and root chord (Cr):
C
 t
(3.24)
Cr
The parameter ftc is a function of thickness ratio and is given by
4
t
t
f tc  1  2.7   100 
 c  max
 c  max
Aircraft Performance Analysis
(3.25)
16
t
where   is the maximum thickness-to-chord ratio of the wing, or the tail. Generally
 c  max
speaking, the maximum thickness to chord ratio for a wing is about 12% to18%, and for a
tail is about 9% to 12%. The parameter Swet in equation 3.12 is the wing or tail wetted
area.
The parameters Swetw, Swetht, and Swetvt denote wetted area of the wing,
horizontal tail, and vertical tail respectively. Unlike reference area, the wetted area is
based on the exposed area, and not the gross area. Due to of the special curvature of the
wing and tail airfoil sections, it may seem time consuming to calculate the accurate
wetted area of a wing or a tail. There is a simplified method to determine the wetted area
of a lifting surface with an acceptable accuracy. Since the wing and tail are not too thick
(average about 15%), it may be initially assumed that, the wetted area is about twice that
of the net or exposed area (see figure 3.9). To be more accurate, you may assume the
lifting as a thing box with average thickness equal to half of the airfoil thickness.
According to this assumption, the wetted area is given as:
Tip Chord
Centerline Chord
Mean Aerodynamic Chord
Root Chord
Figure 3.10. Wing Mean Aerodynamic Chord (MAC)
Aircraft Performance Analysis
17

t 
S wet  21  0.5  bC
 c  max 

(3.26)
For ultimate accuracy, you need to employ a CAD software package (e.g.,
AutoCad or SolidWorks) to calculate it with a high accuracy.
The parameter Cdmin in equations 3.19, 3.20, and 3.21 is the minimum drag
coefficient of the airfoil cross section of the wing or tail. It can be readily extracted from
a Cd-Cl curve of the airfoil. One example is illustrated in figure 3.11 for a six series
NACA airfoil 631-412. Reference [3] is a rich collection of information for a variety of
NACA 4 digits, 5 digits, and 6-serieis airfoils. For instance, the NACA airfoil 631-412
has a minimum drag coefficient of 0.0048 9 for the clean or flap-up configuration.
Example 3.1
Consider a cargo aircraft with the following features:
m = 380,000 kg S = 567 m2, MAC = 9.3 m, (t/c)max = 18%, Cdmin = 0.0052
This aircraft is cruising at sea level with an airspeed of 400 knot. Assume the aircraft CDo
is three times the wing CDo (i.e., CDow), determine the aircraft CDo.
Solution:
Re 
V C 1.225  400  0.5144  9.3

 131,334,640  1.31  108
5

1.785  10
(3.16)
M
V 400  0.5144

 0.605
a
340
(1.20)
Due to the high Reynolds number, the air flow over the wing is turbulent, so
0.455
0.455
Cf 

 0.00205
2.58
log 10 Re
log 10 131,334,6402.58
f M  1  0.08M 1.45  1  0.08  0.606
1.45
 0.9614
(3.15a)
(3.18)
4
t
t
4
f tc  1  2.7   100   1  2.70.18  1000.18  1.591
 c  max
 c  max

t 
S wet  21  0.5  bC  21  0.5  0.18 567  1,236 m 2
 c  max 

C Dow  C f w f tcw
 S wetw
f M 
 S
 C d min w 


 0.004 


Aircraft Performance Analysis
(3.25)
(3.26)
0.4
(3.19)
18
C Do  0.00205  1.591  0.9614 
w
1236  0.0052 


576  0.004 
0.4
 0.0759
Therefore the aircraft zero-lift drag coefficient is:
C Do  2.3C Dow  3  0.0759  0.0228
Cdmin
Figure 3.11. The variations of lift coefficient versus drag coefficient for airfoil NACA 631412 (Ref. 3)
Aircraft Performance Analysis
19
3.4.3. High Lift Devices (HLD)
High lift devices are parts of the wing to increases lift when employed (i.e., deflected).
They are usually employed during take-off and landing. Two main groups of high lift
devices are leading edge high lift devices (often called flap) and trailing edge high lift
devices (e.g., slat). There are many types of wing trailing edge flaps such as split flap,
plain flap, single-slotted flap, fowler flap, double-slotted flap, and triple-slotted flap.
They are deflected down to increase the camber of the wing, in order to increase lift, so
the maximum lift coefficient, CLmax will be increased. The most effective method used on
large transport aircraft is the leading edge slat. A variant on the leading edge slat is a
variable camber slotted Kruger flap used on the Boeing 747. The main effect of the wing
trailing edge flap is to increase the effective angle of attack of the wing without actually
pitching the airplane. The application of high lift devices has a few negative side-effects
including an increase in aircraft drag (as will be included in CDo).
3.4.3.1. Trailing Edge High Lift Device
The increase in CDo due to application of trailing edge high lift device (flap) is given by
the following empirical formula:
Cf 
 A   f B
C Do flap  
(3.27)
 C 
where Cf/C is the ratio between average flap chord to average wing chord (see figure
3.12) at the flap location and is usually about 0.2. In case where the flap is extended, as in
the case for a slotted flap, the Cf represents the extended chord. Furthermore, when a flap
is extended, the wing chord (C) is also locally increased. The equation 3.27 is based on a
flap with a flap-span-to-wing-span ratio of 70%. In the case where the flap span is
different, the results should be revised accordingly.
f
Cf
C
a. Flap; only deflected (Plain flap)
f
Cf
C
b. Flap; extended while deflected (Slotted flap)
Figure 3.12. Wing section at the flap location
Aircraft Performance Analysis
20
Do not confuse this Cf with skin friction coefficient which shares the same
symbol. The parameters “A” and “B” are given in the Table 3.3 based on the type of flap.
The f is the flap deflection in degrees (usually less than 50 degrees).
No
1
2
3
4
5
Flap type
Split flap
Plain flap
Single-slotted flap
Double-slotted flap
Fowler
A
B
0.0014 1.5
0.0016 1.5
0.00018 2
0.0011
1
0.00015 1.5
Table 3.3. The values of A and B for different types of flaps
3.4.3.1. Leading Edge High Lift Device
The increase in CDo due to application of a leading edge high lift device (e.g., slot, and
slat) is given by the following empirical formula:
C 
C Do   sl C Do
sl
 C  w
(3.28)
where Csl/C is the ratio between average extended slat chord to average extended wing
chord. The CDow is the wing zero-lift drag coefficient without extending a high lift device
(including slat).
3.4.4. Landing Gear
The landing gear (or undercarriage) is the structure (usually struts and wheels) that
supports the aircraft weight and facilitates its motion along the surface of the runway
when the aircraft is not airborne. Landing gear usually includes wheels and is equipped
with shock absorbers for solid ground, but some aircraft are equipped with skis for snow
or floats for water, and/or skids. To decrease drag in flight, some landing gears are
retracted into the wings and/or fuselage with wheels or concealed behind doors; this is
called retractable gear. In the case of retracted landing gear, the aircraft clean CDo is not
affected by the landing gear.
When landing gear is fixed (not retracted) in place, it produces an extra drag for
the aircraft. It is sometimes responsible for an increase in the aircraft drag as high as
50%. In some aircraft, a fairing is used to decrease the drag of a non-retracted gear (see
figure 3.14b). The fairing is a partial cover that has a streamlined shape such as airfoil.
The increase in CDo due to wheel of the landing gear is given by the following empirical
equation:
n
 S lg
C Do   C Dlg  i
lg
i 1
 S




Aircraft Performance Analysis
(3.29)
21
where Slg is the frontal area of each wheel, and S is the wing reference area. The
parameter CDlg is the drag coefficient of each wheel; that is 0.15, when it has fairing; and
is 0.30 when it does not have any fairing (see figure 3.13). The frontal area of each wheel
is simply the diameter (dg) times the width (wg).
S lg  d g wg
(3.30)
dg
wg
a. Without fairing
b. With fairing
Figure 3.13. Landing gear and its fairing
As it is observed in equation 3.29, every wheel that is experiencing air flow must be
accounted for drag. For this reason, subscript “i” is used. The parameter “n” is the
number of wheels in an aircraft. The drag calculation for the strut of landing gear is
presented in the next Section. Some aircraft are equipped with skid, especially when they
have tail gear configuration. Skid is not a lifting surface, but for the purpose of zero-lift
drag calculation, it may be treated as a small wing.
3.4.5. Strut
In this section, we deal with two types of strut: 1. Landing gear strut, and 2. wing strut.
Landing gear is often attached to the aircraft structure via strut. In some GA/homebuilt
aircraft, wings are attached through a few struts to support wing structure; i.e., strutbraced (see figure 3.14). Modern aircraft use advanced material for structure that are
stronger and there are no need for any strut to support their wings; i.e. cantilever. In some
aircraft (such as hang gliders), the cross section of the wing strut is a symmetrical airfoil
in order to reduce the strut drag. In both cases, the strut is producing an extra drag for
aircraft.
The increase in CDo due to application of strut is given by the following empirical
equation:
Aircraft Performance Analysis
22
n
S 
CDos   CDosi  s 
S 
i 1
(3.31)
where Ss is the frontal area of each strut (its diameter times its length), and S is the wing
reference area. The parameter CDos is the drag coefficient of each strut; that is 0.1, when it
is faired (i.e., has an airfoil section). When strut does not have an airfoil section, its drag
coefficient is obtained from Table 3.1. The parameter “n” is the number of struts in an
aircraft. It is observed that using an airfoil section for a strut decreases its drag by an
order of 10. However, its manufacturing cost is increased too.
strut
Figure 3.14. Wing strut and landing gear strut in a Cessna 172
3.4.6. Nacelle
If the engine is not buried inside fuselage, it must be in a direct contact with air flow. In
order to reduce the engine drag, engine is often located inside an aerodynamic cover
called nacelle. For the purpose of drag calculation, it can be considered that the nacelle is
similar to the fuselage, except its length-to-diameter ratio is lower. Thus, the nacelle zerolift drag coefficient (CDon) will be determined in the same way as in a fuselage. In the
case where the nacelle length-to-diameter ratio is below 2, assume 2. This parameter is
used in the equation 3.17. Some aero-engine manufacturers publish the engine nacelle
drag in the engine catalog, when the installed drag is demonstrated. In such case, use that
manufacturer data. Determine nacelle drag, only when the engine uninstall thrust is
available.
3.4.7. External Fuel Tank
An external fuel tank (such as a wingtip tank) is in a direct contact with air flow and is
generating drag. An external fuel tank may be modeled as a small fuselage. In the case
where the fuel tank length-to-diameter ratio is below 2, assume 2.
Aircraft Performance Analysis
23
3.4.8. Cooling Drag
An aeroengine is a source of heat where it is generated when fuel is burned in the
combustion chamber. This heat is initially conducted to the engine external surface such
including cylinder heads via oil/water. In order to keep the engine efficient and maintain
its performance, the heat should be transferred to outside airflow through a convenient
technique; usually conduction. Various types of heat exchangers are arranged in an
aeroengine which require a flow of air through them for the purposes of cooling. The
source of this cooling air is usually the free air stream possibly augmented to some extent
by a propeller slipstream (in a prop-driven engine) or bleed from the compressor section
(in a turbine engine).
As the air flows through, to extract heat energy from the body, it experiences a loss in the
total pressure and momentum. This loss during heat transfer process may be interpreted
as a drag force referred to as the cooling drag.
Turbine engine manufacturers calculate the net power/thrust lost in the flow and
subtract this from the engine uninstalled power/thrust. In such case, no cooling drag
increment is added to the aircraft. Piston engine manufacturer often does not provide
installed power, so the cooling drag needs to be determined in order to calculate the
installed power. For a piston engine, the engine power is typically reduced by as much as
approximately 5% to account for the cooling losses.
Cooling air flow
Air flow
Cowl flap opened
Exit air
flow
Figure 3.15. The use of cowl flaps to control engine cooling air
An engine needs to be installed to the aircraft structure (e.g., fuselage and wing)
through a special mounting/pylon. An air cooled engine has a special mounting (See
figure 3.15) to provide the airflow around the engine. The calculation of the cooling drag
is highly installation-dependent. It is unfeasible to present a general technique which can
be applied to every engine installation. Turbine/Piston engines manufacturers use
Aircraft Performance Analysis
24
computer software packages for the calculation of installation losses for their engines.
Because of the complexity of the engine configuration and the internal flow through a
typical engine installation, current methods for estimating cooling losses are empirical in
nature.
For an oil-cooled turbine engine, it is sufficient to consider cooling drag by
consulting with the engine manufacturer catalog and using installed power/thrust. For an
air-cooled engine, engine cooling drag coefficient (CDoen) is given [14] by the following
empirical relationship:
CDoen  4.51108
K e PT 2
VS
(3.32)
where P is the engine power (in hp), T is the hot air temperature (in K) at exit,  is the
relative density of the air, V is the aircraft airspeed (in m/sec), and S is the wing reference
area (in m2). The parameter Ke is a coefficient that depends on the type of engine and its
installation. It varies between 1 and 3.
3.4.9. Trim Drag
Basically, trim drag is not basically different from the types of drag already discussed. It
arises mainly as the result of having to produce a horizontal tail load in order to balance
the airplane around its center of gravity. Any drag increment that can be attributed to a
finite lift on the horizontal tail contributes to the trim drag. Such increments mainly
represent changes in the induced drag of the tail.
To examine this further, we begin with the sum of the lifts developed by the wing
and tail that must be equal with the aircraft weight in a trimmed cruising flight.
Lt  Lw  L  W
(3.33)
where
1
(3.34)
Lw  V 2 S w C Lw
2
1
(3.35)
Lt  V 2 S t C Lt
2
Equation 3.33 is one of the requirements of the aircraft longitudinal trim requirements in
cruising flight. Substituting tail and wing lift into equation 3.33, one can derive the
horizontal tail lift coefficient (CLt) as:
S
(3.36)
C Lt  C L  C Lw
St


where CL is aircraft lift coefficient, CLw is wing lift coefficient, and St is horizontal tail
area. Then, the trim drag will be:
C Dotrim  C Dit
St
St 
1
S
 KtC

 C L  C Lw

S   et  ARt S 
St 
2
Lt
Aircraft Performance Analysis


2
(3.37)
25
where et is horizontal tail span Oswald efficiency factor, and ARt is the horizontal tail
aspect ratio. The trim drag is usually small, amounting to only 1% or 2% of the total drag
of an airplane for the cruise condition. Reference 5, for example, lists the trim drag for
the Learjet Model 25 as being only 1.5% of the total drag for the cruise condition.
3.4.10. CDo of Other Parts and Components
So far, we introduced techniques to calculate CDo of aircraft major components. There are
other components, parts, factors, and items that are producing drag and contribute to total
the CDo of an aircraft. These items are introduced in this subsection.
1. Interference
When two bodies intersect or are placed in proximity, their pressure distributions and
boundary layers can interact with each other, resulting in a net drag of the combination
that is often higher than the sum of the separate drags. This increment in the drag is
known as interference drag. Except for specific cases where data are available,
interference drag is difficult to calculate accurately. Four examples are: 1. placing an
engine nacelle in proximity to a rear pylon (e.g., Gulfstream V), and 2. the interference
drag between the wing and fuselage in a mid-wing configuration. 3. horizontal tail aft of
jet engine exhaust nozzle, and 4. wing leading edge aft of a propeller wake.
a. High wing
c. Low wing
Interference
b. Mid wing
d. Parasol wing
Figure 3.16. Wing-fuselage interference drag
Figure 3.16 shows a wing attached to the side of a fuselage. At the fuselage-wing
juncture a drag increment results as the boundary layers from the two components
interact and thicken locally at the junction. This type of drag penalty will become more
severe if surfaces meet at an angle other than 90o. Reference [4] shows that the
interference drag of a strut attached to a body doubles as the angle decreases from 90o to
approximately 60°. Avoid acute angles, if not, filleting should be applied at the juncture
to reduce interference drag.
In the case of a high-wing configuration, interference drag results principally from
the interaction of the fuselage boundary layer with that from the wing’s lower surface.
This latter layer is relatively thin at positive angles of attack. In addition, it is the
boundary layer on the upper surface of a low wing that interferes with the fuselage
boundary layer. The upper surface boundary layer is relatively thicker than the lower
Aircraft Performance Analysis
26
surface boundary layer. Thus, the wing-fuselage interference drag for a low-wing aircraft
is usually greater than for a high-wing aircraft.
The available data on interference drag are limited; References [2] and [4]
presents a limited amount. Based on this reference, an approximate drag increment
caused by wing-fuselage interference is estimated to equal 4% of the wing’s zero-lift drag
for a typical aspect ratio and wing thickness.
Theoretical accurate calculation of interference drag is impossible, unless you
utilize a high fidelity computational fluid dynamics (CFD) software package. For
example, a wing protruding from a fuselage just forward of the station where the fuselage
begins to taper may trigger a flow separation over the rear portion of the fuselage.
Sometimes interference drag can be favorable as, for example, when one body operates in
the wake of another. The most accurate technique to determine interference drag is to
employ wind tunnel or flight test.
2. Antenna
The communication antennas of many aircraft are located outside of the aircraft. They are
in direct contact with the air, hence they produce drag. There are mainly two types of
antennas: a. rod or wire, and b. blade or plate. Both type of antennas should be in an open
space (below on above the fuselage) for a better communication. During a flight,
airstream flows around an antenna, so it will generate drag.
Figure 3.17. The UHF radio antenna of a reconnaissance aircraft U-2S Dragon Lady
Figure 3.18. Antenna of Boeing Vertol CH-113 Labrador
Aircraft Performance Analysis
27
Figure 3.17 shows a rod antenna of reconnaissance aircraft U-2S Dragon Lady.
The large plate antenna of a reconnaissance aircraft U-2S is the UHF radio antenna. The
thin whip acts as an ADF antenna. The whip was originally straight. The bend in the
upper portion of the whip antenna was introduced to provide clearance for the senior
span/spur dorsal pod. Most reconnaissance aircraft U-2 seem to now use the bent whip,
even if they are not capable of carrying the senior span/spur dorsal pod.
Figure 3.18 illustrates lower fuselage section of a Boeing Vertol CH-113
Labrador with various antennas and plates. Note the two blade antennas, and two whip
antennas. Also seen, is the towel-rack 'loran' antenna. The item hanging down, in the
distance, is the sectioned flat plate that covers the ramp hinge, when the ramp is closed.
a. pitot tube mounted under the wing of a Cessna 172
b. Boeing 767 fuselage nose with three pitot tubes [16]
Figure 3.19. pitot tube
Aircraft Performance Analysis
28
To calculate CDo of antenna, rod antenna is treated as a strut, and blade antenna is
considered as a small wing. The very equations that are introduced for strut (Section
3.4.5) and wing (Section 3.4.2) may be employed to determine the CDo of an antenna.
3. Pitot Tube
A pitot tube (combined with a static port) is a pressure measurement instrument used to
measure airspeed and sometimes altitude of an aircraft (see figure 3.19). The basic pitot
tube simply consists of a tube pointing directly into the air flow. As this tube contains air,
a pressure can be measured as the moving air is brought to rest. This pressure is
stagnation pressure of the air, also known as the total pressure, or sometimes the pitot
pressure.
Since the pitot tube is in direct contact with the outside air, it has a contribution to
aircraft drag (i.e., CDo). If the aircraft is in subsonic regime, the horizontal part of the pitot
tube may be assumed as a little fuselage and its vertical part as a strut. In a supersonic
flight, shock wave is formed around a pitot tube, and creates additional wave drag. For
supersonic flow, section 3.5 introduces a technique to account for pitot tube zero-lift drag
coefficient (CDo).
4. Surface Roughness
The outer surface of an aircraft structure (skin) has a considerable role in aircraft drag.
For this reason, the aircraft skin is often finished and painted. The paint not only protects
the skin from atmospheric hazards (e.g., rusting), but also reduces its drag. As the surface
of the skin is more polished, the aircraft drag (surface friction) will be reduced. The
interested reader is referred to specific aerodynamic references and published journal
papers to obtain more information about the effect of surface roughness on the aircraft
drag.
5. Leakage
There are usually gaps between control surfaces (such as elevator, aileron and rudder),
flaps, and spoilers and the lifting surfaces (such as wing and tails). The air is flowing
through these tiny gaps and thus producing extra drag. This is called leakage drag.
Leakage drag is usually contributing about 1% - 2% to total drag. The reader is referred
to specific aerodynamic references and published journal papers to gain more information
about the influence of these gaps on the aircraft drag.
6. Rivet and Screw
The outside of the aircraft structure (fuselage and wing) is covered with a skin. The skin;
if it is from metallic materials; is attached to the primary components of an aircraft
structure (such as spar and frame) via parts such as rivets or screws. The Composite
structures have the advantage that they do not require any rivet or screw, and are often
attached together through techniques such as bonding.
In the case of the screw, the top part of the screw could be often hidden inside
skin. But the heads of most types of rivets are out of skin, therefore they produce extra
drag. Figure 3.19a illustrates both rivet and screw on wing of Cessna-172. Rivets and
Aircraft Performance Analysis
29
screws are usually contributing about 2% -3% to total drag. The author did not find any
reference which demonstrates the calculation of rivet contribution to aircraft drag. you
may use flight test to measure this type of drag.
7. Pylon
A podded turbine engine (e.g., turbofan engine) is usually mounted to the airplane wing
or fuselage with a pylon. Pylon is mainly a structural component, but may be designed
such that it contributes to aircraft lift. Although pylon is not fundamentally a lifting
surface, but for the purpose of zero-lift drag calculation, it may be treated as a small
wing.
8. Compressibility
Compressibility effects on aircraft drag have been important since the advent of
high-speed aircraft in the 1940s. Compressibility is a property of the fluid, when a fluid is
compressed, its density will be increased. Liquids have very low values of
compressibility whereas gases have high compressibility. In real life, every flow of every
fluid is compressible to some greater or lesser extent. A truly constant density
(incompressible) fluid is a myth. However, for almost all liquids (e.g., water), the density
changes are so small that the assumption of constant density can be made with reasonable
accuracy.
0.02
0.018
0.016
CDo
CDo
O
*
x
0.014
0.012
0.01
0.008
0.9
0.95
1
M
1.05
1.1
Mach number
Figure 3.20. Drag rise due to compressibility for swept wing-body combinations (Ref. 1)
As a fluid is compressible, a flow of fluid may be compressible too.
Incompressible flow is a flow where the density is assumed to be constant throughout.
Compressible flow is a flow in which the density is varying. A few important examples
are the internal flows through rocket and gas turbine engines, high-speed subsonic,
transonic, supersonic, and hypersonic wind tunnels, the external flow over modern
airplanes designed to cruise faster than Mach 0.3, and the flow inside the common
Aircraft Performance Analysis
30
internal combustion reciprocating engine. For most practical problems, if the density
changes by more than 5 percent, the flow is considered to be compressible. Figure 3.20
shows the effect of compressibility on drag for three configurations.
Flow velocities higher than Mach 0.3 are associated with a considerable pressure
changes, accompanied by correspondingly noticeable changes in density. The aircraft
drag at high subsonic speed (e.g., M = 0.9) is about twice as of that at low subsonic speed
(e.g., Mach 0.2). Consider a wing with a free stream. The lift is created by the occurrence
of pressure higher than free stream on the lower surface of the wing and lower than free
stream on the upper surface. The is usually coincides by the occurrence of velocities
higher than free stream on the upper surface of the wing and lower than free stream on
much of the lower surface.
As the airspeed approaches the speed of sound (i.e., M > 0.85), due to a positive
camber, the higher local velocities on the upper surface of the wing may reach and even
substantially exceed local sonic speed, M = 1 (See Figure 3.21). The free stream Mach
number at which the maximum local Mach number reaches 1 is referred to as the critical
Mach number.
M = 1 (sonic line)
M = 0.85
M>1
Shock wave
M<1
Figure 3.21. Local supersonic speed with a high subsonic flow
The existence of sonic and supersonic local velocities on a wing is associated with
an increase in drag. The extra drag is due to a reduction in total pressure through oblique
shock wave which causes local adverse pressure gradients. In addition, when an oblique
shock interacts with boundary layer, the boundary layer gets thicker and may even
separates. This is referred to as compressibility drag.
A cambered airfoil has typically a crest and a lifting surface with such airfoil has
a crestline. The crest is the point on the airfoil upper surface to which the free stream is
tangent. The crestline is the locus of airfoil crests along the wing span. The drag increase
due to compressibility is generally not large, until the local speed of sound occurs at or
behind the crest of the airfoil. Empirically, it is found for all airfoils except the
supercritical airfoil, that at 2% to 4% higher Mach number than that at which M = 1 at the
crest the drag rises abruptly. The Mach number at which this abrupt drag rise begins is
called the Drag Divergence Mach number.
Aircraft Performance Analysis
31
-3
x 10
7
6
5
CDc
CDc
4
3
2
1
0.8
0.85
0.9
0.95
1
1.05
M
Ratio of free stream Mach number to crest critical Mach number
Figure 3.22. Incremental drag coefficient due to compressibility (Ref. 1)
However, the occurrence of substantial supersonic local velocities well ahead of
the crest does not lead to a significant drag increase provided that the velocities decrease
again below sonic speed forward of the crest. The incremental drag coefficient due to
compressibility is designated CDc. Figure 3.22 is an empirical average of existing
transport aircraft data. Table 3.4 demonstrates the incremental drag coefficient due to
compressibility for several aircraft.
No
Aircraft
1
2
3
4
Boeing 727
North American F-86 Sabre
Grumman F-14 Tomcat
McDonnell Douglas F-4
Phantom
Convair F-106 Delta Dart
0.013
0.02
54%
Table 3.4. Drag increase due to compressibility effects for several aircraft
5
CDo at low
subsonic
speed
0.017
0.014
0.021
0.022
CDo at high
subsonic
speed
0.03
0.04
0.029
0.031
Incremental drag
coefficient due to
compressibility (CDc)
76.5%
207.7%
38.1%
40.9%
A complete set of drag curves for a typical large transport jet is given in Figure 3.23.
9. Miscellaneous items
There are other items which produce drag, but due to their variations, were not listed in
this section. These include such items as 1. footstep in small GA aircraft which allow the
pilot or technician to climb up; 2. small plate on tails to balance the manufacturing
unbalances, 3. catapult hook, 4. pipe and funnel for aerial fuel refueling, 5. wiper blades,
6. angle of attack meter vane, 7. non-streamlined fuselage, 8. deflector cable, 9. chemical
Aircraft Performance Analysis
32
sprayer in agricultural aircraft, 10. vortex generator, 11. navigation lights, and 12. fitting.
The calculation of drag of such minor items is out of scope of this text.
0.04
M=0.1
M=0.2
M=0.3
M=0.4
M=0.5
M=0.6
0.035
CD
0.03
CD
0.025
0.02
0.015
0.5
0.55
0.6
0.65
0.7
M
0.75
0.8
0.85
0.9
Mach number
Figure 3.23. Drag coefficient versus Mach number
3.4.10. Overall CDo
Based on the build-up technique, the overall CDo is determined readily as the sum of CDo
of all aircraft components and factors. The calculation of CDo contribution due to factors
as introduced in section 3.4.9 is complicated. These factors sometimes are responsible for
an increase in CDo up to about 50%. Thus we will resort to a correction factor in our
calculation technique. Then, the overall CDo of an aircraft is given by:


C Do  K c C Dow  C Do f  C Doht  C Dovt  C Dos  C Dolg  C Don  C Do ft  ...
No
1
2
3
4
5
6
7
8
Aircraft type
Jet transport
Agriculture
Prop Driven Cargo
Single engine piston
General Aviation
Fighter
Glider
Remote Controlled
(3.38)
Kc
1.1
1.5
1.2
1.3
1.2
1.1
1.05
1.2
Table 3.5. Correction factor (Kc) for equation 3.38
where Kc is a correction factor and depends on several factors such as the type, year of
fabrication, degree of streamline-ness of fuselage, configuration of the aircraft, and
number of miscellaneous items. Table 3.5 yields the Kc for several types of aircraft.
Aircraft Performance Analysis
33
Each component has a contribution to overall CDo of an aircraft. Their
contributions vary from aircraft to aircraft and from configuration to configuration (e.g.,
cruise to take-off). Table 3.6 illustrates [2] contributions of all major components of
Gates Learjet 25 to aircraft CDo. Note that the row 8 of this table shows the contributions
of other components as high as 20%.
No
1
2
3
4
5
6
7
8
9
Component
Wing
Fuselage
Wing tip tank
Nacelle
Engine strut
Horizontal tail
Vertical tail
Other components
Total CDo
CDo of component
0.0053
0.0063
0.0021
0.0012
0.0003
0.0016
0.0011
0.0046
0.0226
Percent from total CDo (%)
23.4
27.8
9.3
5.3
1.3
7.1
4.8
20.4
100
Table 3.6. CDo of major components of Gates Learjet 25
No Aircraft
Type
1
Cessna 172
2
Cessna 182
3
Cessna 185
4
Cessna 310
5
6
7
8
Learjet 25
Gulfstream II
Saab 340
Single
engine GA
Single
engine GA
Single
engine GA
Twin
engine GA
Business jet
Business jet
Transport
Airliner
9
10
11
12
Boeing 707-320
Boeing 767-200
Airbus 340-200
Boeing C-17
Globemaster III
McDonnell
Douglas D-C-9-30
Airliner
Airliner
Airliner
Transport
Landing
gear
Fixed
mTO
(kg)
1,111
S
(m2)
16.2
Engine
CDo
Piston
0.028
Fixed
1,406
16.2
Piston
0.029
Fixed
1,520
16.2
Piston
0.021
Piston
0.026
2 Turbojet
2 Turbofan
2 Turboprop
2 Turbofan
0.0226
0.023
0.028
0.021
Retractable 2,087
Retractable
Retractable
Retractable
Retractable
6,804
29,711
29,000
49,090
21.53
75.21
41.81
92.97
Retractable
Retractable
Retractable
Retractable
151,320
4 Turbofan
142,880 283.3 2 Turbofan
275,000 361.6 4 Turbofan
265,350 353
4 Turbofan
0.013
0.0135
0.0165
0.0175
Table 3.7. CDo of several aircraft at low speed
Table 3.7 illustrates CDo of several aircraft at low speeds [15]. The retired airliner Boeing
707 was a very aerodynamic aircraft with a CDo of 0.013. You may have noticed that
Cessna 185 has a lower CDo compared with Cessna 172 and Cessna 182. The reason is
Aircraft Performance Analysis
34
that Cessna 172 and Cessna 182 are equipped with fixed tricycle landing gear while
Cessna 185 has a fixed tail dragger.
3.5. Wave Drag
In supersonic airspeeds, a new type of drag is produced and it is referred to as “shock
wave drag” or simply “wave drag”. When a supersonic flow experiences an obstacle
(e.g., wing leading edge or fuselage nose); it is turned into itself and a shock wave is
formed. A shock wave is a thin sheet of air across which abrupt changes occur in flow
parameters such as pressure, temperature, density, speed, and Mach number. In general,
air flowing through a shock wave experiences a jump toward higher density, higher
pressure, higher temperature, and lower Mach number. The effective Mach number
approaching the shock wave is the Mach number of the component of velocity normal to
the shock wave. This component Mach number must be greater than 1.0 for a shock to
exist. The wave drag; the new source of drag; is inherently related to the loss of the
stagnation pressure and increase of entropy across the oblique and normal shock waves.
In general, a shock wave is always required to bring supersonic flow back to
subsonic regime. In a subsonic free stream, whenever the local Mach number becomes
greater than 1 over the surface of a wing or body, the flow must be decelerated to a
subsonic speed before reaching the trailing edge. If the surface could be shaped such that
the surface Mach number is reduced to 1 and then decelerated subsonically to reach the
trailing edge at the surrounding free stream pressure, there would be no shock wave and
no shock drag. A major goal of transonic airfoil design is to reduce the local supersonic
Mach number to as close to 1 as possible before the shock wave. One of the main
functions of sweep angle in a swept wing is to reduce wave drag at transonic and
supersonic airspeeds.
In supersonic airspeeds, three main types of waves are created: 1. oblique shock
wave, 2. normal shock wave, and 3. Expansion waves. In general, an oblique shock wave
brings a supersonic flow to another supersonic flow with a lower Mach number. A
normal shock wave brings a supersonic flow to a subsonic flow (i.e., Mach number less
than 1). However, an expansion wave brings a supersonic flow to another supersonic
flow with a higher Mach number. In supersonic airspeeds, a shock wave may happen at
any place in the aircraft, including wing leading edge, horizontal tail leading edge,
vertical tail leading edge, and engine inlet. All these aircraft components locations will
create an extra drag when a shock wave is formed. Thus, in supersonic speed, the drag
coefficient is expressed by:
CD  CDo  CDi  CDw
(3.39)
where CDw is referred to as "wave drag coefficient". The precise calculation of C Dw is
time consuming, but to give the reader the guidance, we present two techniques, one for a
lifting surface leading edge; and one for the whole aircraft. For a complicated geometry
aircraft configuration, the drag may be computed with aerodynamics techniques such as
vortex panel method and “Computational Fluid Dynamics” techniques. Figure 3.24
illustrates Lockheed Martin F-35 Lightning Joint Strike Fighter, a new generation of
Aircraft Performance Analysis
35
fighters. It is claimed that F-35 will be the last manned fighters, and in the near future, the
fighters will be unmanned.
Figure 3.24. Lockheed Martin F-35 Lightning Joint Strike Fighter, a new generation
of fighters
3.5.1. Wave Drag for Wing and Tail
In this section, the wave drag of a wing is presented. In an aircraft with a supersonic
maximum speed, the locations where have the first impact with airflow (such as wing
leading edge, horizontal tail leading edge, vertical tail leading edge, fuselage nose, and
engine inlet edge) are usually made sharp. One main reason for this initiative is to reduce
the number of shack waves created over the aircraft components during a supersonic
flight. Two airfoil cross sections which are frequently employed for wing, horizontal tail,
and vertical tail are double wedge and biconvex (See figure 3.25).
a. double wedge
b. biconvex
Figure 3.25. Supersonic airfoil sections
In this section, a wave drag calculation technique is introduced which is applicable to any
component where has a corner angle (e.g., double wedge and biconvex) and experiences
a shock wave. Then wave drag coefficient (CDw) for such component is calculated
separately and then all CDw are summed together.
Consider a the front top half of a wedge with a sharp corner that experiences an oblique
shock in a supersonic flow (as illustrated in Figure 3.26). The wedge has a wedge angle
of , and the free stream pressure, and Mach number are P1 and M1 respectively. When
Aircraft Performance Analysis
36
the supersonic flow hits the corner, an oblique shock with angle of  to the free stream is
generated.
For each component with such configuration, the wave drag coefficient is given by:
Dw
Dw
C Dw 

(3.40)
1
1
2
2
V S
  M PS
2
2
where subscript infinity (  ) means that the parameter are considered in the infinity
distance from the surface. Please note that, this does not really mean infinity, but it
simply means at a distance out of the effect of shock and the surface (i.e., free stream).
The parameter S is the surface of a body at which the pressure is acting (i.e., planform
area).
Dw is the wave drag and is equal to the axial component of the pressure force. In
supersonic speeds, the induced drag may be ignored, since it has negligible contribution,
compared to the wave drag. In supersonic speeds, the lift coefficient has a very low value.
Since, the wedge has an angle, ; the drag force due to flow pressure actin on a wedge
surface with an area of “A” is given by:
Dw  P2 Asin  
(3.41)
where P2 represents the pressure behind an oblique shock wave, and  is the corner angle
(see figure 3.26). Please note that; in this particular case, the pressure at the lower surface
is not changed. The relationship between upstream and downstream flow parameters may
be derived using energy law, mass conservation law, and momentum equation. The
relationship between upstream pressure (P1) and downstream pressure (P2) is given [7]
by:
oblique
shock
M2
P2
P1
M1


DW
Figure 3.26. Geometry for drag wave over a wedge


2
P2  P1 1 
M n21  1 
(3.42)
  1

where  is the is the ratio of specific heats at constant pressure and constant volume ( =
cp/cv). For air at standard condition,  is 1.4. The variable Mn1 is the normal component of
the upstream Mach number, and is given by:


Aircraft Performance Analysis
37
M n1  M 1 sin 
(3.43)
The parameter  oblique shock angle is a nonlinear function of upstream Mach
number and wedge corner angle ():
 M 12 sin 2   1 
tan   2 cot   2
(3.44)

 M 1   cos 2   2 
Equation 3.44 is called -M relation. For any given , there are two values of 
predicted for any given Mach number. Take the lower value which is the representation
of the weal shock wave which is favored by the nature.
The pressure may also be non-dimensionalized as follows:
P2  P 
1
1
 V 2 C p    M  2 P C p
2
2
(3.45)
where Cp is the pressure coefficient.
Now, consider another more general case, where a supersonic airfoil has an angle of
attack, or when the second half of a wedge is considered. In such a case, where a
supersonic flow is turned away from itself, an expansion wave is formed (See figure
3.27). Expansion waves are the antithesis of shock waves. In an expansion corner, the
flow Mach number is increased, but the static pressure, air density, and temperature
decrease through an expansion wave. The expansion wave is also referred to as the
Prandtl-Meyer expansion wave.
Mach lines
Oblique shock wave
M2, P2
M1, P1


M3, P3
C
Figure 3.27. Oblique shock and Prandtl-Meyer expansion waves
For a wing with a wedge airfoil section shown in figure 3.27, the drag force is the
horizontal components of the pressure force, so:
(3.46)
D  P2 A sin   P3 A sin 
where A is the surface area, and is equal to:
Aircraft Performance Analysis
38
Ab
C 2
cos 
(3.47)
where b is the wing span (not shown in the figure). Plugging the equation 3.46 into
equation 3.47, the following result is obtained:
C 2
1
(3.48)
D  P2  P3 b
sin   P2  P3 bC tan 
cos 
2
If the wing has a different geometry, use this basic technique to determine the wing wave
drag. The pressure behind the expansion wave is given [7] by the following equation:

 1
 1

2 
1  2 M 2 
(3.49)
P3  P2 

1    1 M 32 


2
where the flow Mach number behind expansion wave (M3) and the flow Mach number
behind expansion wave (M2) are related through the turn angle,  :
(3.50)
   M 3   M 2 
The (M) is referred to as the Prandtl-Meyer function and is given by:

  1 1    1 2
(3.51)
 (M ) 
tan 
M  1   tan 1 M 2  1
 1
  1

The equation 3.51 should be employed twice in the calculation; once to calculate M2, and
once to determine M3. Example 3.3 indicates the application of the technique.




Example 3.2:
M1 = 2
P1
y
P2
10o
D

x
P1
P2
C
Figure 3.28. Geometry for Example 3.2 (side-view)
Consider a rectangular wing with a span; b (not shown in the figure) of 5 m and a chord;
C of 2 m in a Mach 2 air flow (the side-view is sketched in figure 3.28). The airfoil
Aircraft Performance Analysis
39
section is a 10° half-angle wedge at zero angle of attack. Calculate the wave drag
coefficient. Assume that the expansion over the corners of the base is such that, the base
pressure is equal to the free-stream pressure. The experiment is performed at sea level
ISA condition.
Solution:
At sea level ISA condition, the free-stream pressure is 1 atm (i.e., 101,325 Pa), and air
density is 1.225 kg/m3. The ratio of specific heats at constant pressure and constant
volume for air,  is 1.4. Since the supersonic flow turns to itself, an oblique shock wave is
produced both at top and bottom surfaces. Pressure is increased behind the shock wave
(P2). In order to determine the pressure P2, we need to first calculate the shock wave
angle ():

2 2 sin 2   1 
 M 12 sin 2   1 
tan   2 cot   2
 (3.44)
  tan(10)  2 cot   2
 2 1.4  cos 2   2 
 M 1   cos 2   2 
The solution of this nonlinear equation yields an angle of 39.3 degrees ( = 39.3 deg).
The normal component to the shock Mn1 is:
(3.43)
M n1  M 1 sin   2 sin39.3  1.27
The static pressure behind the shock wave is:


2
 2(1.4)



P2  P1 1 
M n21  1  101,3251 
1.27 2  1
 1.4  1

  1

(3.42)
 P2  172,364 Pa
The wave drag is the net force in the x direction; P2 is exerted perpendicular to the top
and bottom surfaces. The force due to P1 exerted over the base is in the direction opposite
to the x axis. S is the planform area (the projected area seen by viewing the wedge from
the top, thus S = c × b. The drag, Dw, is the summation of two top and bottom pressure
force (see Figure 3.29) components in x direction, minus the force from base pressure:

 P2 S 
 cos  


A
S
cos 

 P2 S 
 cos   sin 


S
Figure 3.29. Wave drag, the component of pressure force in x direction
Dw  2P2 A2 sin    P1 A1
Aircraft Performance Analysis
(3.41)
40
where A2 is the base area, and A1 is the top and bottom surfaces:
A1  2bC tan(10)
C
A2  b
cos(10)
Thus:
 C 
Dw  2 P2 b
sin 10  P1b2C tan10
 cos 10 
Dw  2  172,364  5  2  tan10  2  110,325  5  2  tan10  251,670 N
By definition, the wave drag coefficient is:
Dw
251,670
C Dw 

2
1
  M  2 P S 0.5  1.4  101,325  2  5  2
2
(3.40)
 C Dw  0.089
Note that this drag coefficient is based on the shock wave, and it includes the contribution
of the base.
Example 3.3:
Consider a very thin flat plate wing with a chord of 2 m and a span of 5 m. The wing is
placed at a 5o angle of attack in a Mach 2.5 air flow (Figure 3.30). Determine the wave
drag coefficient. Assume sea level ISA condition.
M1 = 2.5
P1
y
5o
x
P2
P3
Figure 3.30. Geometry for Example 3.3 (side-view)
Solution:
a. In the lower surface, an oblique shock is formed. The pressure P3 is determined as
follows:
Aircraft Performance Analysis
41
 2.5 2 sin 2   1 
 M 12 sin 2   1 
(3.44)
tan   2 cot   2

  tan(5)  2 cot   2
 2 1.4  cos 2   2 
 M 1   cos 2   2 
The solution of this nonlinear equation yields an angle of 27.4 degrees ( = 27.4 deg).
The normal component to the shock Mn1 is:
(3.43)
M n1  M 1 sin   2 sin27.4  1.15
The static pressure behind the oblique shock wave is:


2
 2(1.4)



P3  P1 1 
M n21  1  101,3251 
1.15 2  1
 1.4  1

  1

(3.42)
 P3  139,823 Pa
b. In the upper surface, an oblique shock is formed. The pressure P2 is determined as
follows:
For forward Mach line:

  1 1    1
2
2
 1 (M ) 
tan 
M 1  1   tan 1  M 1  1 


 1
  1


 1 (M ) 


 1.4  1

1.4  1
tan 1 
2.5 2  1   tan 1
1.4  1
 1.4  1




(3.51)
 2.5  1  39.1 deg
2
   M 2   M 1    M 2    M 1     39.1  5  44.12 deg
For rearward Mach line:

  1 1    1
2
2
 2 (M ) 
tan 
M 2  1   tan 1  M 2  1 


 1


1



44.12 






(3.51)
(3.50)
(3.51)

 1.4  1

1.4  1
2
2
tan 1 
M 2  1   tan 1  M 2  1 



1.4  1
 1.4  1

(3.51)
Solution of this nonlinear equation yields:
M 2  2.723
The static pressure behind the expansion wave is:

(3.51)
1.4
1.4  1
   1 2   1

2  1.4 1


1

M
1

2
.
5
1




2
2
P2  P1 
 71,742 Pa
  101325

1    1 M 22 
1  1.4  1 2.7232 




2
2
Aircraft Performance Analysis
(3.49)
42
c. Drag
The wave drag force is the x-component of the pressure force:
Dw  P3  P2 bC sin   139,823  71,512  5  2  sin5  59,336 N
(3.41)
The wave drag coefficient is:
C Dw 
Dw
1
  M 1 2 P1 S
2
59,336
0.5  1.4  2.5 2  101,325  5  2

(3.40)
 C Dw  0.0134
3.5.2. Aircraft Wave Drag
An aircraft has usually a complex geometry and various components. In previous section,
mathematical technique was introduced to calculate the wave drag where the object has a
very simple (e.g., flat plate, wedge) configuration. Every aircraft component contributes
to wave drag. In this section, a technique is presented to determine the wave drag of a
complete aircraft.
In this approach, we consider an aircraft as a whole and we will not divide it into several
components. This is an approximate and empirical technique [8]. The aircraft wave drag
coefficient consists of two parts; volume-dependent wave drag (CDwv), and lift-dependent
wave drag (CDwl). The volume-dependent wave drag is a function of aircraft volume and
much greater than lift-dependent wave drag. The reason is that, at supersonic speeds, the
lift coefficient (CL) is very minimal.
(3.52)
C Dw  C Dw  C Dw
l
v
The lift-dependent wave drag is given by:
K SC 2 M 2  1
CDwl  wl L 2
(3.53)
2L
where L represents the aircraft fuselage length, S is wing reference area, and K wl is a
parameter given by


2
 S 
K wl  2 
 bL 
and b is the wing span.
Aircraft Performance Analysis
(3.54)
43
Mach number
Figure 3.31. Wave drag for several aircraft (Ref. 9)
The volume-dependent wave drag is given by:
128K wvV 2
SL4
where V is the total aircraft volume and Kwv is a factor given by:
b

 1  0.75 
L
K wv  1.17
b
 1  2



L 

and  is a function of Mach number as
CDwv 
  M 2 1
(3.55)
(3.56)
(3.57)
In general, wave drag is significant, such that it will increase the aircraft drag up
to about two to three times compared with subsonic drag. For instance, CDo of fighter
aircraft F-4 at Mach number 1.5 (supersonic) is about 0.045, while its CDo at Mach
number 0.6 (subsonic) is about 0.016. Figure 3.31 shows zero-lift drag coefficients of
several aircraft for a range of Mach numbers from subsonic to supersonic speeds. The
drag rise of most of the aircraft due to high Mach number is considerably high. Table 3.8
illustrates CDo of several aircraft at their cruise speed.
3.6. CDo for Various Configurations
Any individual aircraft will often take various configurations in flight phases. When an
aircraft retracts its landing gear, deflects its flap, rotates its control surfaces, exposes any
external component (such as gun), releases its store (e.g., missile), or opens its cargo
door; it has changed its configuration. In general, there are three configuration groups
Aircraft Performance Analysis
44
that an aircraft may adopt, they are: 1. clean configuration, 2. take-off configuration, and
3. landing configuration.
The clean configuration is the aircraft configuration employed at flight phases
such as cruise, climb and turn. As the names imply, the take-off configuration, and
landing configuration are the aircraft configurations which are adopted during take-off
and landing respectively. In an aircraft with a retractable landing gear, the landing gear
are employed during take-off and landing and retracted during cruising flight.
Furthermore, the flaps are extended in take-off and landing operations. The flap
deflection angle is a function of aircraft payload weight and atmospheric conditions.
No
Aircraft
Engine
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
Pilatus PC-9
Lockheed F-104 Starfighter
Embraer EMB 312 Tucano
Boeing 747
Lockheed Jetstar
Lockheed C-5 Galaxy
Boeing 727
Grumman F-14 Tomcat
Learjet 25
Douglas C-54 Skymaster
Curtiss C-46 Commando
Beech V35 Bonanza
Cessna 310
McDonnell Douglas F-4 Phantom
Piper PA-28
Cessna 172
North American F-86 Sabre
North American XB-70 Valkyrie
F-106 Convair Delta Dart
McDonnell Douglas (now Boeing)
F/A-18 Hornet
Turboprop
Turbojet
Turboprop
Turbofan
Turbojet
Turbofan
Turbofan
Turbofan
Turbofan
Piston prop
Turboprop
Turboprop
Piston prop
Turbojet
Piston prop
Piston prop
Turbojet
Turbojet
Turbojet
Turbofan
No. of Landing
Engines gear
1
1
1
4
4
4
3
2
2
4
2
1
2
2
1
1
1
6
1
2
Retractable
Retractable
Retractable
Retractable
Retractable
Retractable
Retractable
Retractable
Retractable
Retractable
Retractable
Retractable
Retractable
Retractable
Fixed
Fixed
Retractable
Retractable
Retractable
Retractable
CDo
0.022
0.016
0.021
0.018
0.0185
0.019
0.07
0.02
0.022
0.023
0.025
0.025
0.025
0.03
0.047
0.028
0.014
0.006
0.013
0.017
Table 3.8. CDo of several aircraft
3.6.1. Clean Configuration
The clean configuration is the configuration of an aircraft when it is at a cruise flight
condition. At this configuration, no flap is deflected, and landing gear is retracted (if it is
retractable). Therefore the drag polar is
C Dclean  C Do
clean
 
 K C LC
2
(3.58)
Thus, clean CDo of the aircraft ( CDo
clean
) includes every component (such as wing, tail, and
fuselage), and excludes flap and landing gear (if retractable). If the landing gear is not
Aircraft Performance Analysis
45
retractable (e.g., as in Cessna 172), the CDo includes landing gear too. The parameter CLC
is the cruise lift coefficient.
3.6.2. Take-Off Configuration
The take-off configuration is the configuration of an aircraft when it is at take-off
condition. In this configuration, the aircraft has high angle of attack, flap is deflected for
take-off, and landing gear is not retracted (even if it is retractable). In a take-off
condition, the flaps are usually deflected down about 10-30 degrees. The take-off CDo
depends on type and the deflection angle of the flaps. As the flap deflection increases, the
take-off CDo increases too. The drag polar at take-off configuration is
2
CDTO  CDo  K CLTO
(3.59)

TO

where, the take-off zero-lift drag coefficient is given by
C DoTO  C Doclean  C DoflapTO  C DoLG
where C Doflap 
TO
(3.60)
represents the zero-lift drag coefficient of flap at take-off condition. In
addition, the parameter C Do represents the zero-lift drag coefficient of the landing gear.
LG
Moreover, CLTO represents the lift coefficient at take-off. This coefficient does not have a
constant value during take-off, due to the accelerated nature of the motion. The CLTO at
the lift-off condition (where the front wheel is just detached from the ground), is given by
2mg
(3.61)
S VLO 2
where VLO represents the aircraft lift-off speed. The factor of 0.9 is added due to the
contribution of the engine thrust during take-off in vertical direction. The aircraft lift-off
speed is often about 10%-30% higher than the aircraft stall speed.
VLO  K LOVS
(3.62)
CLTO  0.9
where KLO = 0.1 - 0.3. In Chapter 8, the take-off performance will be discussed in great
details.
3.6.3. Landing Configuration
The landing configuration is the configuration of an aircraft when it is at landing
condition. In this configuration, the aircraft has high angle of attack (even more than
take-off condition), flap is deflected (even more than take-off condition), and landing
gear is not retracted (even if it is retractable). In landing condition, the flaps are usually
deflected down about 30-60 degrees. The landing CDo depends of the deflection angle of
the flaps. As this angle increases, the landing CDo increases too. The landing zero-lift drag
coefficient (CDoL) is often greater than the take-off zero-lift drag coefficient (CDoTO). If
there is another means of high lift device for the aircraft, such as slat, you need to add it
to this equation. The drag polar at landing configuration is given by:
2
(3.63)
C D  C D  K C L 
L
oL
L
Aircraft Performance Analysis
46
where C LL is the lift coefficient at landing and is given by:
CLL 
2mg
S VL 2
(3.64)
where VL is the aircraft landing speed. The landing speed (VL) and is often about
10%-30% greater than the stall speed.
VL  K LVS
where KL = 0.1 - 0.3. The landing zero-lift drag coefficient is given by
(3.65)
CDoL  CDoclean  CDoflapL  CDoLG
(3.66)
where CDoflap represents the zero-lift drag coefficient of the flap at landing condition.
L
Both CDoflap
TO
and CDoflap are functions of flap deflection. In Chapter 8, the landing
L
performance will be presented.
0.05
M=0.3
M=0.7
M=0.8
M=0.9
0.045
0.04
CD
0.035
CD
0.03
0.025
0.02
0.015
0
0.1
0.2
0.3
CLCL
0.4
0.5
0.6
0.7
Figure 3.32. Drag coefficient versus lift coefficient for various Mach number
3.6.4. The Effect of Speed and Altitude on CDo
The aircraft zero-lift drag coefficient is a function of a number of flight parameters.
Reynolds number is one of the influential parameters on the zero-lift drag coefficient. As
the Reynolds number increases, the boundary layer thickness decreases and thus CDo
decreases as well. As equation 3.16 indicates, the Reynolds number is a function of true
airspeed. Since the true airspeed is a function of altitude (indeed, air density), it can be
concluded that the Reynolds number is also a function of altitude. Another factor
Aircraft Performance Analysis
47
affecting CDo is the compressibility which is significant at speeds higher than Mach 0.7.
The third important factor is the wave drag as discussed in section 3.5. Considering these
factors, it is concluded that the CDo is a function of Mach number and altitude:
C Do  f M , h
(3.67)
At low subsonic Mach numbers, CDo is increased, due to an increase in Reynolds
number. As compressibility factor shows up in higher subsonic speeds, the CDo increases
with a higher rates. In transonic speeds, shock wave is formed and a jump (sudden
increase) in CDo will be experienced. Therefore, the CDo is directly proportional with
speed, and as speed (Mach number) is increased, the CDo increases. Figure 3.32 illustrates
typical variations of drag coefficient versus lift coefficient at various subsonic Mach
numbers.
The second factor that affects the CDo is altitude. For a specific Mach number, as
the altitude increases, the true airspeed decreases. For instance, consider an aircraft is
flying with a speed of Mach 0.5 at sea level. The true airspeed at this altitude is 170
m/sec (0.5 × 340 = 170). If this aircraft is flying with the same Mach number at 11,000 ft
altitude, the true airspeed will be 147 m/sec (0.5 × 249 = 147). In addition, the air density
decreases with altitude with a higher rate. Thus, the higher altitude means the lower
Reynolds number (Equation 3.22), and therefore the higher CDo. Figure 3.33 illustrates
the variations of drag force for a light transport aircraft with turbofan engines at various
altitudes. This transport aircraft has a stall speed of 90 knots and a maximum speed of
590 knots.
5
2
x 10
h=0
h=2000 m
h=4000 m
h=6000 m
h=8000 m
h=10000 m
h=12000 m
1.8
1.6
D (N)
1.4
1.2
1
0.8
0.6
0.4
0.2
0
100
200
300
V (knot)
400
500
600
Figure 3.33. Variations of drag force without considering the compressibility effects
Aircraft Performance Analysis
48
In conclusion, it can be assumed that at the Mach numbers less than 0.7, the
variation of CDo is not significant such that it can be considered constant. At higher Mach
numbers, the compressibility effect and wave drag must be taken into account. The
second conclusion is that, at higher altitude, total drag force is reduced. The reason is
that, although at higher altitude, the CDo is increased, but the air density decreases. The
rate of change (decrease) in air density is faster than the rate of change (increase) of CDo.
This is one of the reasons why airlines choose to fly at higher altitudes despite the need
and cost of climb.
Example 3.4
Consider the aircraft in example 3.1 is equipped with a single-slotted flap with an average
flap chord of 2.3 m. The aircraft takes off with a flap angle of 20 degrees and lands with a
flap angle of 35 degrees. Assume the CDo of landing gear is 0.01, induced drag correction
factor; K is 0.052; and both take-off and landing speeds are 130 knot. Determine the
aircraft CDo at take-off and landing configurations.
Solution:
The CDo of the flap is given by
Cf 
 A   f B
C Do flap  
 C 
From Table 3.3, A = 0.00018 and B = 2, so
(3.26)
The CDo of flap at take-off is
 2.3 
2
C Do flap  
  0.00018  20  0.0178
 9.3 
(3.26)
The CDo of the flap at landing is
 2.3 
2
C Do flap  
  0.00018  35  0.0545
 9.3 
The take-off CDo:
(3.26)
C DoTO  C Doclean  C DoflapTO  C DoLG  0.023  0.0178  0.01  0.051
(3.47)
To find aircraft drag coefficient, we need to determine the induced drag coefficient. The
lift coefficient at take-off is:
C LTO  0.9
2mg
S VLO 
2

0.9  2  380000  9.81
1.225  567  130  0.5144
2
C Di  KC L  0.052  2.16 2  0.242
Aircraft Performance Analysis
2
 2.16
(3.49)
(3.12)
49
So, CDTO is

C DTO  C Do  K C LTO
TO

2
 0.051  0.243  0.293
For the landing:
2mg
2  380000  9.81
C LL 

 2.4
2
S VLO  1.225  567  130  0.51442
2
(3.48)
(3.52)
C Di  K  C L  0.052  2.4 2  0.3
(3.12)
CDoL  CDoclean  CDoflapL  CDoLG  0.023  0.0545  0.01  0.088
(3.47)
 
C DL  C DoL  K C LL
2
 0.088  0.3  0.387
Aircraft Performance Analysis
(3.48)
50
Problems
Note: In all problems, assume ISA condition, unless otherwise stated.
1. A GA aircraft is flying at 5,000 ft altitude. The length of the fuselage is 7 m, wing
mean aerodynamic chord is 1.5 m, horizontal tail mean aerodynamic chord is 1.5
m, and vertical tail mean aerodynamic chord is 0.6 m. Determine Reynolds
number of fuselage, wing, horizontal tail and vertical tail.
2. The following (figure 3.34) is a top-view of Boeing 757 transport aircraft that has
a wing span of 38.05 m. Using a proper scale and using a series of measurements,
determine the wing reference (gross) area, and wing net area of this aircraft.
Figure 3.34. Top-view of Boeing 757 transport aircraft
3. Estimate the wing wetted area of Boeing 757 (problem 2). Assume the wing has a
maximum thickness of 12%.
4. The mean aerodynamic chord of a trainer aircraft is 3.1 m. This trainer is cruising
at sea level with a speed of Mach 0.3. Determine skin friction coefficient of the
wing when boundary layer over the wing is: a. laminar, b. turbulent.
5. A business jet aircraft with a 31 m2 wing area and a mass of 6,500 kg is flying at
10,000 ft altitude with a speed of 274 knot. If C Do = 0.026, K = 0.052, CLmax =
1.8, calculate and plot the following items:
Aircraft Performance Analysis
51
a. drag polar
b. variation of drag force with speed.
6. The tip chord of a wing is 6 m and its root chord is 9 m. What is the mean
aerodynamic chord?
7. The Attack aircraft Thunderbolt II (Fairchild A-10) has the following features:
mTO = 22,221 kg, S = 47 m2, K = 0.06, CDo = 0.032, Vmax = 377 knot, VTO = 120
knot. Assume that the CDo is constant throughout all speeds.
a. Plot the Variation of zero-lift drag (Do) versus speed (from take-off speed
to maximum speed)
b. Plot the Variation of induced drag (Di) versus speed (from take-off speed
to maximum speed)
c. Plot the Variation of total drag (D) versus speed (from take-off speed to
maximum speed)
d. At what speed, the drag force is minimum?
8. The wing of the twin-turbofan airliner Boeing 777 has a 31.6 degrees of leading
edge sweepback, a span of 60.93 m and a planform area of 427.8 m2. Determine
Oswald efficiency factor (e) and induced drag correction factor (K) for this wing.
9. Determine the Oswald efficiency factor of a rectangular wing with an aspect ratio
of 14.
10. A single engine trainer aircraft with a wing area of 26 m2 has a fixed landing gear
with three similar tires. Each tire has a diameter of 25 cm and a thickness of 7 cm.
Each tire is connected to the fuselage with a strut with a diameter of 4 cm and a
length of 15 cm. The landing gear does not have fairing, determine zero-lift drag
coefficient of the landing gear.
11. A cargo airplane is cruising with a speed of Mach 0.47, is taking off with a speed
of 95 knot and is landing with a speed of 88 knot. The plain flap with a 20%
chord ratio is deflected down 22 degrees during take-off and 35 degrees during
landing. The aircraft has a mass of 13,150 kg, a wing area of 41.2 m2, and induced
drag correction factor, K is 0.048. The zero lift drag coefficients of all
components are
CDow = 0.008, CDof = 0.006, CDoht = 0.0016, CDovt = 0.0012, CDon = 0.002,
CDoLG = 0.015, CDos = 0.004,
Determine drag force at
a. cruise
b. take-off
c. landing
12. A Swedish aircraft designer is designing a fuselage for a 36-passenger aircraft to
cruise at a Mach number of 0.55. He is thinking of two seating arrangement
Aircraft Performance Analysis
52
options: a. 12 rows of three passengers, and b. 18 rows of two passengers. If he
selects option a, the fuselage length would be 19.7 m with a diameter of 2.3 m. In
option b the fuselage length would be 27.2 m with a diameter of 1.55 m. What
option yields the lowest fuselage zero-lift drag coefficient?
13. The amphibian airplane Lake LA-250 (figure 3.35) has a wing with the following
features:
S = 15.24 m2, b = 11.68 m, MAC =1.35 m, (t/c)max = 15%, airfoil: NACA 4415,
Cdminw = 0.0042
The aircraft has a mass of 1,678 kg and is cruising with a speed of 155 knot.
Determine the wing zero-lift drag coefficient. For other information, use the
aircraft three-view that is provided.
Figure 3.35. Three-view of amphibian airplane Lake LA-250
14. Assume that the aircraft in problem 13 has a plain flap with a chord ratio of 0.2.
Determine the flap CDo, when it is deflected 30 degrees during a take-off
operation.
15. A twelve-seat commuter airplane has a vertical tail with the following features:
Svt = 4.3 m2, bvt = 5.4 m, (t/c)max_vt = 10%, Cdmin_vt = 0.005. The aircraft has a
wing area of 29.4 m2 and is cruising with a true airspeed of 160 knot at an altitude
of 18,000 ft ISA condition. Determine the vertical tail zero-lift drag coefficient.
16. The four-seat light airplane Piper PA-34 has a horizontal tail with the following
features:
Sht = 3.6 m2, bht = 4.14 m, (t/c)max_ht = 12%, Cdmin-ht = 0.0047
The aircraft has a mass of 2,154 kg, a wing area of 19.39 m2 and is cruising at
171 knot at an altitude of 18,500 ft. Determine the horizontal tail zero-lift drag
coefficient.
17. A jet trainer with a retractable landing gear has the following features:
Aircraft Performance Analysis
53
mTO = 5,630 kg, S = 25.1 m2, b = 17.4 m, e = 0.85, Vc = 270 knot, CLmax = 2.2, VTO =
1.2 Vs, VLand = 1.3 Vs, CDoclean = 0.032, CDoflapTO = 0.02, CDoflapLand = 0.035, CDoLG =
0.01.
Determine the aircraft drag at three flight conditions: a. clean (cruise speed) , take-off,
and c. landing.
18. The wings of the Y-5B agricultural biplane are connected together thru two struts.
Each strut has a circular cross section with the length of 1.2 m and a diameter of 4
cm. The total planform area of both wings is 38 m2; determine the zero-lift drag
coefficient of struts.
19. A single-engine small aircraft has a wing area of 26 m2 and a turboprop engine
that has a power of 600 hp. The aircraft is cruising at 25,000 ft altitude with a
speed of 182 knot. Determine the engine cooling drag coefficient. Assume Ke = 2.
20. A business jet aircraft with a wing area of 45 m2 is equipped with two turbofan
engines. Each engine is podded in a nacelle with a diameter of 62 cm and a length
of 110 cm. Determine nacelle zero-lift drag coefficient, provided the aircraft is
cruising with a speed of 320 knot at an altitude of 28,000 ft.
21. A maneuverable aircraft has two fuel tanks at the wingtips. Each fuel tank has a
shape of an ellipsoid with a diameter of 20 cm and a length of 130 cm. Determine
fuel tank zero-lift drag coefficient, provided the aircraft is cruising with a speed of
190 knot at an altitude of 18,000 ft.
22. The F-16 supersonic jet fighter (figure 3.36) with a mass 12,331 kg and a wing
area of 27.8 m2 is cruising at an altitude of 40,000 ft at Mach number 2.1. The
wing span is 9.45 m and length of the fuselage is 15.3 m. Determine wave drag of
the aircraft, assuming the aircraft volume is 21.3 m2.
23. Assume that F-16 (problem 22) is flying at a Mach number of 1.2 at the same
altitude. What is wave drag of this fighter at this flight condition? Compare the
result with the result of problem 22 and comment about your finding.
Aircraft Performance Analysis
54
Figure 3.36. The three-view of Fighter jet F-16A
24. The fuselage of a large transport aircraft has a diameter of 3.8 m and a length of
43 m. Determine the fuselage zero-lift drag, assuming the wing area to be 180 m2
and cruising speed to be 350 knot at 35,000 ft.
25. Determine the terminal velocity of a water droplet; assuming it has a shape of a
sphere with a radius of 2 mm. The density of water is 998 kg/m3.
26. You are required to design a parachute for the safe landing of a paratrooper with a
mass of 85 kg. Determine the diameter of the parachute, if the terminal velocity is
required to be 1 m/s at 5000 ft altitude. Ignore the drag of the paratrooper and
strings. The mass of the parachute is 6 kg.
27. Felix Baumgartner, an Austrian daredevil fell to Earth from more than 24 miles
high, becoming the first human to break the sound barrier under his own power,
on October 14, 2012. He made the highest and fastest jump in history after
ascending by a helium balloon to an altitude of 128,100 feet, reaching a maximum
speed of Mach 1.24. Assume the terminal velocity was at 70,000 ft, and his
frontal area while descending is 0.12 m2; determine drag coefficient of his body.
28. Consider a very thin flat plate wing with a chord of 0.4 m and a span of 3.6 m.
The wing is placed at a 6o angle of attack in a Mach 1.3 air flow. Determine the
wave drag coefficient.
29. Consider a 2% thickness flat plate wing with a chord of 1.2 m and a span of 4.5
m. The wing is placed at a 3o angle of attack in a Mach 1.7 air flow. Determine
Aircraft Performance Analysis
55
a. wing zero-lift drag coefficient
b. induced drag coefficient
c. wing wave drag coefficient.
Assuming the aircraft has a mass of 1000 kg and is cruising at an altitude of
12,000 ft. The wing Cdmin may be assumed to be 0.003, and Oswald efficiency
factor to be 0.8.
30. Consider a rectangular wing with a span of 15 m and a chord of 3 m in a Mach 2.6
air flow (the side-view as sketched in figure 3.28). The airfoil section is a 7° halfangle wedge at zero angle of attack. Calculate the wave drag coefficient. Assume
that the expansion over the corners of the base is such that, the base pressure is
equal to the free-stream pressure.
31. Consider a tapered wing with a span of 12 m and a chord of 2.5 m in a Mach 1.5
air flow (the side-view as sketched in figure 3.28). The airfoil section is a 4° halfangle wedge at zero angle of attack. Calculate the wave drag coefficient. Assume
that the expansion over the corners of the base is such that, the base pressure is
equal to the free-stream pressure. The wing taper ratio is 0.4.
Aircraft Performance Analysis
56
References
1. Hoak D. E., USAF Stability and Control Datcom, Air Force Flight Dynamics
Laboratory, Wright-Patterson Air Force Base, Ohio, 1978
2. McCormick B. W., Aerodynamics; Aeronautics; and Flight Dynamics, John Wiley,
1995
3. Abbott I. H., and von Doenhoff A. E., Theory of Wing Sections, Dover, 1959
4. Horner S. R., Fluid-Dynamic Drag, Midland Park, NJ, 1965
5. Ross R., Neal R.D., Learjet Model 25 Drag Analysis, Proceeding of the NASAIndustry-University GA drag reduction workshop, Lawrence, KS, July 14-16, 1975
6. Shevell R., Fundamentals of Flight, Second Edition, Prentice Hall, 1989
7. Anderson J., Modern Compressible Flow, Third edition, McGraw-Hill, 2005
8. Kuchemann D., Aerodynamic Design of Aircraft, Pergamon Press, 1978
9. Raymer D. P, Aircraft Design: A Conceptual Approach, AIAA, 2006
10. Anderson J., Fundamentals of Aerodynamics, third edition, McGraw-Hill, 2005
11. Roskam J., Airplane Design, Volume VI, DAR Corp., 2004
12. Paul Jackson, et al., Jane’s all the World’s Aircraft, Jane’s Information Group, several
years
13. Cengel Y. A., and Cimbala J. M., Fluid Mechanics: Fundamentals and Applications,
3rd edition, McGraw-Hill, 2013.
14. Stinton D., The Anatomy of the Airplane, Second Edition, AIAA, 1998
15. Lan E. C. and Roskam J., Airplane Aerodynamics and Performance, DAR Corp.,
2003
16. www.airliners.net
Aircraft Performance Analysis
57